Additives comprising fractions of a percent to several percent of solid propellant mixtures have been considered through the years and are commonly employed in many rocket propellants and explosives. Various additives include burn-rate modifiers (e.g., ferric oxide, metal oxides, and organometallics); curing agents; and plasticizers. In certain cases, additions of small (<5% by weight) amounts of powdered material to the propellant mixture have been shown to increase or otherwise favorably modify the burn rate as described in T B Brill, B T Budenz 2000 “Flash Pyrolysis of Ammonium Percholrate-Hydroxyl-Terminated-Polybutadiene Mixtures Including Selected Additives,” Solid Propellant Chemistry, Combustion, and Motor Interior Ballistics, Vol. 185, Progress in Astronautics and Aeronautics, V Yang, T Brill, W-Z Ren (Ed.), AIAA, Reston, Va.: 3-32. For example, it has been observed by a few investigators that TiO2 (titania) particles may enhance stability by creating burn rates that are insensitive to pressure over certain pressure ranges as disclosed in U.S. Pat. No. 5,579,634 issued to Taylor on Dec. 3, 1996. It is suspected that other organometallic particles may produce these and other favorable traits described in Brill. Nanoparticle additives may have an even further influence on the burn rate because of their high surface-to-volume ratios.
Over the past few years, nanoparticles of many different compounds and combinations have received considerable attention in the scientific and engineering research communities. This surge of activity is a result of the many favorable characteristics certain materials and applications exhibit when nanoparticles are involved in some fashion. Benefits are certainly seen in composite Al/AP/HTPB-based solid propellant formulations when the micron-scale metal fuel (i.e., Al) is replaced by nanoscale particles as described in P Lessard, F Beaupré, P Brousseau, 2001 “Burn Rate Studies of Composite Propellants Containing Ultra-Fine Metals,” Energetic Materials—Ignition, Combustion and Detonation, Karlsruhe, Germany; 3-6 Jul. 2002: 88. pp. 1-13 and in A Dokhan, E W Price, J M Seitzman, R K Sigman, “Combustion Mechanisms of Bimodal and Ultra-Fine Aluminum in AP Solid Propellant,” AIAA Paper 2002-4173, July 2002. However, little research has been done on the effect of nanosized additives such as organometallics and related burn rate-enhancing and smoke-reducing compounds.
Other prior art made of record includes U.S. Pat. No. 6,503,350 issued to Martin on Jan. 7, 2003, describes propellants such as may be used in solid rocket motors. In one preferred embodiment, the propellant comprises one high energy propellant composition comprising a homogeneous mixture of fuel and oxidizer having a predetermined fuel/oxidizer ratio, wherein individual fuel particles are generally uniformly distributed throughout a matrix of oxidizer, and a low energy propellant comprising a fuel and oxidizer. The amounts of the two propellants are present in amounts which achieve a preselected burn rate.
U.S. Pat. No. 6,605,167 issued to Blomquist on Aug. 12, 2003, discloses an autoignition material that includes a plurality of agglomerates. Each agglomerate comprises an oxidizer material particle. A plurality of metal fuel particles are disposed on the oxidizer material particle. The metal fuel particles are present in a weight ratio effective to chemically balance the oxidizer material particle. The metal fuel particles exothermically react with the oxidizer material particle when the autoignition material is exposed to a temperature of about 80.degree. C. to about 250.degree. C. A thin binder film adheres the metal fuel particles to the oxidizer material particle and maintains the metal fuel particles in intimate contact with the oxidizer particles.
U.S. Pat. No. 6,270,836 issued to Holman on Aug. 7, 2001, describes sol-gel preparation of particles. The gel-coated microcapsules have improved mechanical stress- and flame-resistance. A method for making the gel coated microcapsules is also provided. Phase change materials can be included in the microcapsules to provide thermal control in a wide variety of environments.
U.S. Pat. No. 6,086,692 issued to Hawkins, et al. on Jul. 11, 2000, describes an advanced design for high pressure, high performance solid propellant rocket motors and describes a solid rocket propellant formulation with a burn rate slope of less than about 0.15 ips/psi over a substantial portion of a pressure range and a temperature sensitivity of less than about 0.15%/.degree F. A high performance solid propellant rocket motor including the solid rocket propellant formulation is also provided. The solid rocket propellant formulation can be cast in a grain pattern such that an all-boost thrust profile is achieved.
U.S. Pat. No. 4,881,994 issued to Rudy, et al. Nov. 21, 1989, discloses ferric oxide as burn rate catalyst and use of isocyanate curing agent. The patent describes a method of making a ferric oxide burning rate catalyst that results in a highly active, finely divided burning rate enhancing catalyst. The ferric oxide burning rate catalyst is particularly adapted for use in a composite solid rocket propellant. This process provides an ultra pure, highly active, finely divided burning rate catalyst.
U.S. Pat. No. 4,658,578 issued to Shaw, et al. on Apr. 21, 1987, discloses improved igniter compositions for rocket motors are provided which, when cured, are non-volatile and are capable of igniting under vacuum conditions and burning steadily at reduced pressures.
U.S. Pat. No. 4,655,858 issued to Sayles on Apr. 7, 1987, describes metal/oxidant agglomerates for enhancement of propellant burning rate are prepared from a finely divided metal (aluminum, boron, titanium, etc.), ammonium perchlorate, and a small quantity of the same binder material that goes into the manufacture of the propellant, such as, hydroxyl-terminated polybutadiene crosslinked with a polyisocyanate.